Persistent vortex generating high regression rate solid fuel grain for a hybrid rocket engine

ABSTRACT

A cylindrically-shaped hybrid rocket engine solid fuel grain defines an axial combustion port. A fuel grain material comprises a compounded blend of thermoplastic fuel and aluminum. The fuel grain comprises fused stack layers, each layer comprising a plurality of fused abutting concentric beaded structures arrayed to define the combustion port; the port exhibits a rifling pattern or rifling inducing geometry along the port wall. When an oxidizer is introduced into the combustion port combustion occurs along the exposed port wall. Each beaded structure defines a geometry that increases the combustion surface area while inducing a vortex flow of oxidizer and fuel gas. As each layer ablates, an abutting layer exhibiting a similar geometry, is revealed, undergoes a gas phase change, and ablates. This process repeats and persists until oxidizer flow is terminated or the fuel grain material is exhausted. The fuel grain may be manufactured by an additive manufacturing process.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a reissue application of U.S. Pat. No. 10,309,346(filed as U.S. patent application Ser. No. 15/867,852 on Jan. 11, 2018),which is a continuation-in-part application of U.S. patent applicationSer. No. 15/262,661, filed on Sep. 12, 2016, entitled Persistent VortexGenerating High Regression Rate Solid Fuel Grain for a Hybrid RocketEngine and Method for Manufacturing Same, now issued U.S. Pat. No.9,890,091, which is a continuation-in-part application of U.S. patentapplication Ser. No. 12/052,934 filed on Mar. 21, 2008 entitled SolidFuel Grain for a Hybrid Propulsion System of a Rocket and Method forManufacturing Same, now issued U.S. Pat. No. 9,453,479, which claimspriority to the provisional patent application No. 60/896,296 filed onMar. 27, 2007 of the same title. The entire disclosure of each of thesedocuments is incorporated herein by reference.

FIELD OF THE INVENTION

The present invention relates to a solid fuel grain for a hybrid rocketengine having a fuel grain port featuring a wall surface pattern orgeometric shape that induces a vortex flow of atomized liquid or gaseousoxidizer as it is urged through the fuel grain port, which dually servesas the rocket engine's combustion chamber, and via an ablation process,the engine's fuel source. Unlike prior art examples, which upon ignitionquickly ablate and dissipate, the present invention's vortex-inducingport wall surface pattern or geometric shape persists as fuel isablated; thus, providing more thorough oxidizer/fuel mixing andcombustion throughout the rocket engine's operation.

BACKGROUND OF THE INVENTION

The present invention relates generally to rocket propulsion systems andspecifically to hybrid rocket engines. There are three basic types ofchemical rockets in use today: liquid rocket engines that use liquidpropellants (also referred to as either liquid monopropellant or liquidbi-propellent engines), solid rocket motors that use solid propellants,and hybrid rocket engines that use a combination of liquid and solidpropellants.

In a conventionally designed hybrid rocket engine, the fuel is stored inthe solid state, while the oxidizer is stored in either a liquid orgaseous state. Traditionally in most hybrid rocket engine designs, thesolid fuel is cast-molded, extruded, or in some instances machined intoa cylindrically shaped structure referred to as a fuel grain. The fuelgrain is designed and formed to feature one or more internal passagesrunning through its length. These passages are referred to as ports. Thefuel grain port or ports dually serve as the hybrid rocket engine'scombustion chamber or chambers, and through a gas phase change andablation process, the fuel source.

The fuel grain is conventionally housed within a metal orfiber-reinforced polymer composite motor case designed to withstand thepressures and elevated temperatures created during the combustionprocess. The motor case may also feature an internal liner made from ahigh-temperature material to create a thermal barrier to prevent damageor burn-through during rocket engine operation.

The motor case, with fuel grain installed, is attached to a forward captypically machined or cast from high-temperature metal alloys. Theforward cap forms at least a portion of the pre-combustion chamber andhouses the oxidizer injector and ignition system. The aft end of themotor case is attached to an assembly that forms at least a portion ofthe post combustion chamber and allows secure attachment to the rocketnozzle. The assembled motor case with fuel grain installed, forward cap,and aft assembly with attached nozzle is conventionally referred to asthe motor or solid section of the hybrid rocket engine.

In a conventionally designed hybrid rocket engine, liquid or gaseousoxidizer is stored separately in an integrally formed pressure vessel ortank forward of the motor section within the rocket powered vehicle.However, in some designs, liquid or gaseous oxidizer may be storedadjacent to the motor section or even remotely on the vehicle.Conventionally, the tank or pressure vessel stored liquid or gaseousoxidizer is urged through a specially designed plumbing system,typically including a flow control valve to feed oxidizer through one ormore oxidizer injectors housed within the motor section forward cap; andin turn, through the fuel grain port or ports.

The motive force needed to urge the liquid or gaseous oxidizer throughthe oxidizer injector or injectors into the fuel grain port or portswith sufficient flow rate to support combustion may be generated by anyone of several techniques such as the use of a self-pressurizingliquifying gas, enabling an oxygen rich liquid to gas phase change usinga catalyst to cause an exothermic reaction, employing a mechanical boostpump, pre-pressurizing the oxidizer tank with an externally suppliedinert gas, or using an on-vehicle high pressure tank filled with aninert gas to boost oxidizer tank pressure.

Regardless of the configuration or type of liquid or gaseous oxidizerused, the assembly of oxidizer tank, pressurizing system and associatedplumbing is typically referred to as the oxidizer section. Collectively,the motor section and the oxidizer section are referred to as a hybridrocket engine, sometimes also referred to as a hybrid rocket motor.

Hybrid rocket engines offer certain advantages over both solid rocketmotors and liquid rocket engines alike. For example, once ignited, asolid rocket motor cannot be stopped until its propellant is exhausted,and it cannot be throttled or restarted. Hybrid rocket engines, likeliquid rocket engines, can be designed for on-command thrusttermination, throttling, and engine restart. Most liquid monopropellantrocket engines use highly toxic, environmentally damaging propellantsthat are now considered too dangerous and to environmentally unsafe forcontinued use.

Compared to most liquid bi-propellant rocket engines, hybrid rocketengines are significantly less mechanically complex, and therefore morereliable and less expensive to develop, manufacture, and operate. Hybridrockets are ideally suited to use propellants that areself-pressurizing, non-toxic, environmentally benign, operate at ambienttemperatures, and require no specialized equipment for handling,transporting, and loading. Furthermore, hybrid rocket engines, due tothe storage of their propellants in different states of matter, areinherently immune to explosion. Immunity to explosion is of greatimportance to rocket-powered vehicle designers and operators. Theirsuperior safety, mechanical simplicity compared to liquid bi-propellantrocket engines, and environmental friendliness all translate to improvedreliability as well as lower development, manufacturing, and operatingcosts.

Despite all of their aforementioned advantages, conventionally designedhybrid rocket engines using cast-molded solid fuels likehydroxyl-terminated polybutadiene (HTPB), a form of synthetic rubberthat has been the most studied hybrid rocket engine fuel to date, arerarely if ever employed for applications requiring vibration free,consistent high performance. Unfortunately, conventionally designedhybrid rocket engines using cast-molded HTPB as well as othercast-molded solid fuels, including paraffin wax, polyamides, andthermoplastics, have not been able to demonstrate the vibration free,consistent, high performance required for most rocket propulsionapplications.

Excessive vibration and inconsistent performance is even more pronouncedwhen higher energetic additives such as aluminum powder have beenblended into solid fuels like HTPB and paraffin wax. All of thesedisadvantages and inefficiencies are attributable to either the solidfuel material selected or to the fuel grain production methods. To fullyunderstand the efficacy and advantages of the present invention, it isimportant to understand these disadvantages in relation to competingrocket propulsion systems as well as their respective causes.

Comparatively poor hybrid rocket engine performance and their oftenunpredictable, even sometimes dangerous nature can be attributed to: 1)low regression rate, i.e., the rate at which the solid fuel is consumedcompared to solid rocket motors, 2) the build-up of adverse harmonicsinducing unacceptable, sometimes dangerous levels of vibration, 3)excessive solid fuel waste compared to other rocket propulsion systems,4) low specific impulse (Isp) compared to most liquid bi-propellantrocket engines, and 5) inconsistent, unpredictable thrust performancethat renders them unusable in clustered (multiple engines per launchvehicle stage or spacecraft) configurations.

1). Low Regression Rate. For a given selection of fuels andoxidizer-to-fuel mass ratios, the thrust generated by a rocket or anytype of reaction engine is approximately proportional to the mass flowrate. In a hybrid rocket engine, mass flow rate is proportional to fuelgrain regression rate. In a classically designed hybrid rocket engine,particularly those using slow burning fuels like HTPB, the burning rateis further limited by the heat transfer from the relatively remote flameto the fuel grain port surface. One of the physical phenomena thatlimits the burning rate is the blocking effect caused by the injectionof vaporizing fuel into the high-velocity oxidizer gas stream. Given thelinear nature of the oxidizer gas stream, oxidizer/fuel vapor mixing andresulting combustion efficiency is a function of the amount of timeavailable for mixing to occur within a classically designed hybridrocket fuel grain port.

Attempts to increase the burning rate by mixing energetic materials likeAlcoa produced Military Grade 44 aluminum powder (Rockledge, Tex.)(average particle size of 44 microns) with traditional hybrid rocketfuels using cast-molding production methods have been only marginallysuccessful in improving rocket engine performance. Aluminum powder ishighly reactive with oxygen and water. To passivate the material to bestable in atmospheric conditions for safe handling, processing, storing,transporting, and use in a rocket engine, the aluminum particle isallowed to form an outer layer of aluminum oxide (alumina), anon-combustible material that when burned acts as a heat sink causing aloss of temperature and energy within the center port. Factoring in anallowance for an appropriately sized oxidizer tank and associatedplumbing in tandem with a long-thin fuel grain, hybrid rocket enginesare considered ungainly and not a viable option for many applications.

Nano-scale aluminum powder is thought to be the next big advancement inboth solid and hybrid rocketry. Elemental aluminum in nano-scale issignificantly higher in reactivity than micron-scale powder due to itsrelatively high specific surface area. Unfortunately, most attempts tosafely and efficaciously employ this material in both solid and hybridrocketry have not been successful. If allowed to form an alumina shell,effectively consuming a portion of the aluminum core, much of theelemental aluminum's energetic value is lost.

In addition to the challenges associated with obtaining a uniform blendof polymer and metal powder throughout the fuel grain using thecast-molding technique, improved burning rates by use of metal additivessuch as aluminum have only served to exacerbate the problems associatedwith using relatively elastic materials such as HTPB and paraffin waxesas a primary hybrid rocket solid fuel. Moreover, attempts to improve theregression rate further using high energetic material such as ALEXpowder (an ultra-fine aluminum powder produced by the plasma-explosionprocess) have been even less successful and have introduced asignificant potential for spontaneous ignition or explosion stemmingfrom the pyrophoric nature of these ultra-fine powders.

Despite the potential for significant increase in burning rate, on theorder of 30% higher than standard Military grade 44 micron particle sizealuminum powder, employing a material that will spontaneously igniteupon exposure to the atmosphere or explode on contact with water orwater vapor is counter-productive to one of the most significantadvantages of a hybrid rocket engine—its comparative higher safety(i.e., benign failure mode and U.S. Government recognized zero TNTequivalency) compared to other forms of chemical rocketry.

More recent efforts have involved the development of methods tostabilize the nano-scale aluminum particles by encapsulating eachparticle in a polymeric material; thereby, protecting the elementalaluminum from the environment. Some of these approaches, such asimmersion in benzene followed by compounding with styrene to formgranules of aluminum-styrene, have merit and warrant furtherinvestigation. Another developed technique involves a process in whichthe elemental aluminum particle, measuring 15 nm or less, is produced ina reactor simultaneously with the formation of a crystalline polymerouter shell. This passivated material is safe to handle, transport,store, and use as rocket propellant, and the particle core remains 99.9%pure elemental aluminum.

This difference in the combustion scheme of a hybrid rocket enginesignificantly degrades the propellant burning rate compared to a solidrocket motor propellant in which the solid-state oxidizer and fuel arein intimate contact. Consequently, the regression rate, usingconventionally molded fuel grain materials like HTPB is typicallyone-tenth or less than that of most solid rocket propellants.

Structurally soft, HTPB with a Young's Modulus varying between 0.0026GPa and 0.00756 GPa is a common polymeric binder used in solid rocketry.It has been the fuel of choice for over fifty years in many U.S.Government sponsored hybrid rocket propulsion research projects. Most ofthis work has involved integrating multi-port configurations into thefuel grain's design to increase the total fuel grain port surface areaas a means to improve regression rate. Unfortunately, improvements inregression rate using multiport designs have been offset by reduced fuelvolume loading, the build-up of adverse harmonics that induce excessiveand sometimes dangerous levels of vibration, unpredictable thrustperformance, and increased fuel waster. However, excessive vibration,unpredictable thrust performance, and increased fuel waste have alsobeen observed in single port large hybrid rocket engine designs usingboth HTPB as well as faster burning, also structurally soft, paraffinwax with a Young's Modulus of 0.061 GPa. While it is generallyunderstood that regression rate in a hybrid rocket engine is a functionof fuel burn rate and port surface area, the increased regression ratesachieved using multi-port grain configurations have been more thanoffset by reduced reliability, consistency, efficiency, and safety.

2). Adverse Harmonics and Excessive Vibration. In any discussion aboutvibration in a hybrid rocket engine, it is important to keep in mindthat the port within a hybrid rocket fuel grain is the engine'scombustion chamber. Combustion chamber wall integrity is an essentialdesign criterion in any reaction engine. Therefore, it is understandablethat if a combustion chamber wall's structural integrity is degraded orcompromised, chamber performance and reliability would likewise bedegraded or compromised. Logically, an engineer would be reluctant touse a compressible, easily fractured material to fabricate a combustionchamber. But, this is exactly the case when soft, compressible, andfracture prone materials like HTPB and paraffin wax are used toconstruct a hybrid rocket fuel grain and its combustion chamber port orports. To make matters more complex, given the fuel grain is also therocket engine's fuel supply, as fuel is consumed, the port wallcontinually ablates and expands in diameter; thereby, increasingavailable surface area causing an oxidizer-fuel mixture shift fromoxidizer rich to fuel rich combination. According to the prior art, thissituation is addressed by throttling the oxidizer flow.

Materials such as HTPB and paraffin wax are thought to respond to highpressure gases created within the port by compressing the solid fuelagainst the higher-strength motor case; thereby, inducing grainfractures and erosive burning both common occurrences in large scaleHTPB and paraffin wax hybrid rocket engines.

Adverse harmonics exhibited in hybrid rocket engines, particularlypronounced in large-scale variants, is thought to be caused by acompressive-relaxation response by these soft fuels reacting to elevatedchamber pressures, creating a type of trampoline effect. Theseoscillations can build to dangerous vibration levels and even cause acatastrophic over pressurization event. Cast-molded fuel grains madefrom these materials are also prone to structural flaws such as weakspots, air bubbles, hot spots, and fractures that are also known tocause erosive burning and erratic, unpredictable performance. Fuelfragments breaking free and blocking or temporarily blocking therocket's nozzle have also been recorded. These phenomena are consideredeven more problematic in large hybrid rocket engines, especially thoseusing multi-port designs.

3). Excessive Solid Fuel Waste. A certain amount of residual solid fuelis expected in a hybrid rocket engine. However, in a multi-portconfiguration, the amount of non-combusted fuel that is expelled can besignificant and in certain circumstances a safety concern. In multi-portdesigns, as the burn progresses and as fuel is ablated and combusted,the structure between the ports ultimately losses its integrity untilfailure occurs. In these situations, chunks of non-combusted fuel andwebbing material have been known to break free, partially and sometimescompletely blocking the nozzle, which can cause a serious safetyproblem. In multi-port HTPB fueled hybrid rocket engine designs, thetotal amount of residual and unspent fuel can reach 15% or more.

4). Poor Specific Impulse. Expressed in seconds, specific impulse(usually abbreviated Isp) is a measure of the efficiency of rocket andjet engines. By definition, it is the total impulse (or change inmomentum) delivered per unit of propellant consumed and is dimensionallyequivalent to the generated thrust divided by the propellant flow rate.Typically referenced as performance in vacuum for rockets, Isp is aconvenient metric for comparing the efficiency of different rocketengines for launch vehicles and spacecraft.

Generally speaking, there is an inverse relationship between increasedregression rate and Isp in a hybrid rocket. Whereas, regression ratespeaks to the hybrid rocket engine's volumetric efficiency and thrustoutput as a function of fuel grain diameter, Isp relates more to therocket engine's propellant efficiency. Ideally, rocket engine designersattempt to improve both. However, attempts to improve hybrid rocket Isphas mainly focused on evaluating and testing different propellantcombinations. Whereas, a classical hybrid rocket engine uses a liquid orgaseous oxidizer and solid fuel, past experiments have been conducted onengine's that use a solid oxidizer and liquid fuels. While many of theseachieved very high Isp—in the high 300 seconds (in a vacuum), theyproved to be impractical for reasons primarily associated with the needto maintain a hydrocarbon fuel as a solid at cryogenic temperatures.

Other approaches have involved blending energetic materials such asaluminum powder into the fuel grain composition to increase Isp.However, obtaining a consistent, uniform mixture has always been achallenge using cast-molding techniques, especially when moldingmulti-port grains. Most conventionally designed hybrid rocket enginesusing nitrous oxide and polymeric fuel like HTPB average Isp is between270 seconds to 290 seconds (vacuum), the higher figure attained with theaddition of aluminum powder as an additive. While higher than most solidrocket motors, this level of performance is significantly lower thancompeting liquid bi-propellant systems using liquid oxygen andhydrocarbon fuels like kerosene that average between 310-340 seconds.

5). Inconsistent Thrust Performance. Inconsistent, unpredictable thrustin a classical hybrid rocket engine is a direct consequence of all ofthe above listed shortcomings and problems. Inconsistent andunpredictable performance makes it impossible for a hybrid rocket engineto be seriously considered for most rocket propulsion applications anduses. Further, many of the causes of inconsistent thrust performance canbe tied to the cast-molding production process used to fabricate hybridrocket fuel grains. HTPB and paraffin wax fuel grains are typicallycentrifugally cast-molded, with the latter containing a small percentageof polyethylene to improve tensile strength. During the HTPBpolymerizing process, small air bubbles are formed and hot spots arecreated due to incomplete mixing and uneven curing. HTPB fuel grainsrequire up to 90 days or more to fully cure, and even then, theirmaterial characteristics change over time. Small air bubbles are alsoformed during the cooling cycle when fuel grains are cast from paraffinwax. Bubble formation is a function of the shrinkage occurring withinthe wax. In an attempt to reduce or eliminate unwanted air bubbles aswell as other types of grain flaws and hot spots, centrifugal castingmethods, taking up to 120 hours to complete, are routinely employed.Even with these measures, air bubbles, structural cracks, hot spots, andother flaws seem to be chronic for fuel grains made using thecast-molding process.

Therefore, it would be highly desirable to develop a solid fuelpropellant and fuel grain architecture-topology that exhibits: 1)flawless composition, 2) a regression rate approaching that of solidrocket motors, 3) significantly improved thrust consistency, 4) morethorough oxidizer-fuel mixing, 5) greatly improved specific impulse, and6) minimal vibration—all without compromising the many safety,mechanical simplicity, and economic advantages inherent in hybrid rocketpropulsion systems.

SUMMARY OF THE INVENTION

The present invention is a high performance, safe to produce, store,transport, and operate hybrid rocket solid fuel grain made fromthermoplastic or a formulation of thermoplastic and high energetic metaladditives such as aluminum powder.

In one embodiment, the fuel grain can be fabricated using a fuseddeposition type additive manufacturing method and apparatus. The detailsof such an additive fabrication apparatus and process can be found inthe commonly-owned U.S. Pat. No. 9,453,479, entitled Solid Fuel Grainfor a Hybrid Propulsion System of a Rocket and Method for ManufacturingSame, issued on Sep. 27, 2016; and the commonly-owned U.S. Pat. No.9,822,045, entitled Additive Manufactured Thermoplastic-NanocompositeAluminum Hybrid Rocket Fuel Grain and Method of Manufacturing Same,issued on Nov. 21, 2017, and the commonly-owned pending patentapplication Ser. No. 15/818,381, entitled Additive ManufacturedThermoplastic-Nanocomposite Aluminum Hybrid Rocket Fuel Grain and Methodof Manufacturing Same, filed on Nov. 20, 2017 . Each of these patentdocuments is incorporated herein in their entirety.

The invention comprises a solid fuel grain for a hybrid rocket enginehaving a cylindrical shape featuring a center combustion port, comprisedas a stack of fused layers fabricated from a material suitable as ahybrid rocket fuel. Each layer is formed as a plurality of fusedabutting concentric ring-shaped, polygonal or similarly shaped beads ofsolidified material of increasing radii arrayed around the center port.

An oxidizer is introduced into the solid fuel grain through the centerport, with combustion occurring along the exposed surfaces of the solidfuel grain center port wall.

Each concentric ring-shaped bead of fuel grain material possesses anundulating or irregular circular geometry or irregular pattern, theplurality of which forms a surface pattern designed to increase theamount of surface area available for combustion, and when stacked andfused may also form a rifling pattern or geometry that inducesoxidizer-fuel gas vortex flow to improve combustion efficiency.

The fuel grain employs an internal topology that both presents increasedport wall surface area to the flame zone while inducing axialoxidizer-fuel gas flow in a manner that persists during the rocketengine's operation; and depending upon the solid fuel selected, eitherphase changes from solid to a gas or from solid to entrained droplets,and is subsequently ablated.

The solid fuel grain is preferably manufactured using any one of avariety of additive manufacturing machine technologies and techniques,including the techniques described and claimed in the commonly-ownedapplication Ser. No. 15/818,381 and the issued U.S. Pat. Nos. 9,453,479and 9,822,045, as referenced above.

After being loaded into a hybrid rocket engine's solid section,concurrent with ignition actuation to elevate the temperature within thecenter port above the solid fuel's ignition or glass transitiontemperature, a liquid or gaseous oxidizer is introduced into the solidfuel grain through one or multiple injectors along a pathway defined bythe center port causing a thin layer of the center port wall to phasechange.

Depending upon the type of solid fuel used, phase change will occureither from a solid to a gas or from a solid to entrained liquiddroplets along the exposed surface area of the solid fuel grain portwall. The resulting gaseous or entrained liquid fuel then mixes with theoxidizer to form an oxidizer/fuel mixture suitable for hybrid rocketengine combustion. The resulting combusted reaction mass is expelled athigh temperature and pressure through the rocket engine's nozzle(conventional de Laval or aerospike) at supersonic speed to generatethrust.

Each concentric solidified circular, polygonal, or similarlyring-shaped, bead of material possesses a geometry that is designed toexpose more fuel surface along the center port wall for combustion thanwould otherwise be possible if the center port wall were of a smooth,uniform cast-molded design. Starting with the center port wall andworking outward, each beaded concentric ring structure, after undergoingphase change and ablation, is replaced by the next abutting beadedconcentric ring structure. This process is repeated and persiststhroughout the rocket engine's operation until either oxidizer flow isterminated or the solid fuel is exhausted

Unlike prior art constructions that attempt to increase regression rateusing cast-molded multi-port grain architecture featuring smooth portwalls at the sacrifice of fuel loading, increased fuel waste, andinduced excessive vibration, the solid fuel grain of the presentinvention, especially if additively manufactured, supports smooth,consistent rocket engine operation at regression rates previouslyunobtainable in a single port design. Further, by replacing cast-moldingproduction methods with additive manufacturing methods, grain flawschronic to both cast-molded fuel grains made from HTPB and paraffin waxare eliminated.

Another exemplary solid fuel grain suitable for use in a hybrid rocketengine and made in accordance with the present invention is formed asdescribed herein, but with each concentric beaded ring structurepossessing a pattern that both increases the surface area available forcombustion and creates, in its plurality of fused stacked layers, arifling type pattern or geometry within the port wall designed to induceoxidizer swirling flow around the center port axis line rather thanlaminar or streamline flow; thereby, creating a vortex within the centerport to enable oxidizer and gaseous fuel to spend more time within thecenter port to mix and combust more thoroughly than would otherwise bepossible.

Again, as in the above exemplary example, the pattern thus engineeredinto the fuel grain topology will persist throughout the rocket engine'soperation until either oxidizer flow is terminated or the solid fuel isexhausted. Prior art constructions have employed swirling type oxidizerinjectors to induce vortex flow. However, this prior art technique isonly partially effective as it cannot generate axial flow throughout thelength of the fuel grain and its center port.

In prior art embodiments, rifling patterns or rifling inducinggeometries have been imprinted onto the molded fuel grain's port wall asa means to induce axial oxidizer-fuel gas flow. Unfortunately, anyvortex generated in this manner is only momentary due to the surfacepattern being quickly ablated and not repeated.

In contrast, the solid fuel grain of the present invention supportssmooth, consistent rocket engine operation at regression ratespreviously unobtainable in a single port design while improving Isp byenabling more complete oxidizer-fuel mixing and combustion whilesignificantly reducing the amount of wasted fuel. More thoroughcombustion and higher Isp enables hybrid rocket engine designers theopportunity to design hybrid rocket engines with reduced fuel grainaspect ratio as well as with lower oxidizer loading to meet dimensionalrestrictions and performance requirements for rocket propulsionapplications that previously could not consider a hybrid rocket engine.

Using a Acrynotrile Butadiene Styrene (ABS) thermoplastic fuel or a fuelformulation consisting of a blend of ABS thermoplastic and aluminumpowder, the particle size of which can vary from micron to nano scale,the fuel undergoes a phase change from solid to gas vapor along theexposed surface area of the solid fuel grain port wall. The resultingcombined fuel vapor and aluminum powder then mixes with the oxidizer toform a fuel/oxidizer mixture suitable for rocket engine combustion. Theresulting combusted reaction mass is expelled at high temperature andpressure through the rocket engine's nozzle (conventional de Laval oraerospike) at a supersonic speed to generate thrust. Generally, hybridrocket engines using fuel grains consistent with the present inventionfabricated from formulations with a ratio in a range from about 75%ABS/20% aluminum will operate with an Isp (vacuum) ranging from about290 to 300 seconds, while a formulation using as little as 5%nanocomposite aluminum and 95% ABS will operate with an Isp (vacuum) ofabout 320 seconds or higher.

DESCRIPTION OF THE DRAWINGS

FIG. 1 is a perspective view of a solid fuel grain section made inaccordance with the present invention;

FIG. 2 is a top view of the solid fuel grain section of FIG. 1 ;

FIG. 3 is a sectional view of the grain section of FIG. 1 , taken alongline 3-3 of FIG. 2 ;

FIG. 4A is a side view of a solid fuel grain comprised of a plurality ofsolid fuel grain sections;

FIG. 4B is an exploded side view of the solid fuel grain of FIG. 5A;

FIG. 5 is a perspective view of the plurality of solid fuel grainsections of FIG. 45A wrapped with insulating film;

FIG. 6 is a sectional view of an exemplary rocket incorporating thesolid fuel grain of FIGS. 4A, 4B, and 5 ;

FIG. 7 is an enlarged sectional view of the motor case of the rocket ofFIG. 6 , showing a flame configuration; and

FIGS. 8A, 8B, and 8C are top views of the fuel grain section of FIG. 1as successively consumed by a flame.

FIG. 9 depicts the coordinate system and orientation of the fuel grainfor use with FIGS. 10-13 .

FIG. 10 depicts a quarter sectional view of the fuel grain section ofFIG. 1 featuring a concentric corrugation topology grain pattern.

FIGS. 11 and 12 depict quarter sectional views of a fuel grain sectionfeaturing a concentric rifled truncated pyramidal topology grainpattern.

FIGS. 13A and 13B depict a top view and a perspective view of the fuelgrain section of FIG. 1 featuring a concentric rifled polygonal topologygrain pattern.

DETAILED DESCRIPTION OF THE INVENTION

The present invention is a solid fuel grain for a hybrid rocket engine.The solid fuel grain can be manufactured 65 using a fused depositiontype additive manufacturing apparatus as described and claimed in thecommonly-owned application Ser. No. 15/818,381 and the issued U.S. Pat.Nos. 9,453,479 and 9,822,045, as referenced above.

FIGS. 1-3 illustrate various views of an exemplary solid fuel grainsection 10 suitable for use in a hybrid rocket engine. Additionaldetails of fuel grains, including their fabrication, are described andclaimed in one or more of the following commonly-owned and relatedapplications, the contents of which are hereby incorporated byreference:

U.S. Pat. No. 9,453,479 issued on Sep. 27, 2016

U.S. Pat. No. 9,822,045 issued on Nov. 21, 2017

patent application Ser. No. 15/818,381 filed on Nov. 20, 2017

patent application Ser. No. 15/262,661 filed on Sep. 12, 2016

With reference to FIGS. 1-3 , the fuel grain section 10 has a generallycylindrical shape and defines a center port 16. In this exemplaryembodiment, the center port 16 has a substantially circularcross-section, but the center port 16 could have other geometries, suchas a star, clover leaf, or polygon without departing from the spirit orscope of the present invention.

The solid fuel grain section 10 is formed as a fusion (bonded) stack oflayers with each such layer formed as a series of abutting fusedconcentric ring-shaped beads of solidified material 12 arrayed aroundthe center port 16. In one embodiment, a heat gun with an ABS stick isused to bond the individual layers. Viscous ABS is applied to thesectional end caps before aligning and joining the grain sections. As isknown by those skilled in the art, other adhesives can be used to jointhe grain sections.

As is further described below, when incorporated into a hybrid rocketengine, an oxidizer is introduced into the solid fuel grain section 10along a pathway defined by the center port 16, with combustion occurringalong the exposed surfaces (also referred to as the boundary wall orcombustion port wall) of the solid fuel grain section 10 port wall.Accordingly, each concentric ring-shaped structure possesses a geometricpattern 14 that serves to increase the surface area for combustion,compared to a smooth concentric circular structure or smooth walls asconsistent with cast-molded constructions. As each such concentricring-shaped bead ablates or undergoes phase change from either solid togas or solid to entrained liquid droplet, the abutting concentric beadis exposed to the flame sheet. This process continues and persistsduring the hybrid rocket engine's operation until either oxidizer flowis terminated or the solid fuel is exhausted. Unlike prior artconstructions that improve regression rate by increasing the surfacearea exposed to the flame sheet using a multi-port architecture at thesacrifice of fuel loading, the solid fuel grain of the present inventionpresents increased surface area as a means to improve regression rate,but without the disadvantages associated with multi-port configurations.

Although the fuel grain section 10 may be manufactured in various sizesor dimensions, in an exemplary embodiment, the fuel grain section 10 hasan outer diameter, d2, of 19.0 inches. Although a wide range ofdiameters and fuel grain lengths (or sectional lengths) are possible,the center port 16 has an initial diameter, d1, (i.e., beforecombustion) of 4.0 inches in this exemplary embodiment (although alarger diameter is shown in FIG. 3 to enable a better view of theinterior of the fuel grain section 10). Although a fuel grain with anygrain diameter can be fabricated, traditionally a ratio of about 5:1(outer diameter to inner diameter) is used for a hybrid rocket fuelgrain.

Each of the stacked fused layers in this exemplary embodiment would havean approximate thickness ranging from 0.005 inches to 0.015 inchesdepending upon the fabrication technique employed.

In one fabrication technique, each of the stacked layers 12 is formed bythe deposition of viscous polymer which is extruded following a roughlycircular tool path forming a plurality of solidified abuttingring-shaped beads of material. Viewed in cross section as depicted inFIG. 11 , each ring-shaped bead of solidified material 90 is oval orelliptical in cross sectional shape, which flattens on its bottom underits own weight as the material cools and flattens on the top as theweight of the next extruded layer of abutting ring-shaped beads ofmaterial is deposed above it.

As for the concentric ring-shaped beaded structures, the objective is toincrease the surface area presented to the flame zone for combustionwithin the center port 16 in a manner that is persistent throughout thehybrid rocket engine operation. In this exemplary embodiment, and asillustrated in FIGS. 1-3 , the surface pattern presented to the flamezone is characterized by a series of projections and depressions(according to other embodiments the surface pattern comprises aplurality of ribs, a plurality of undulations, a plurality ofprotrusions and recesses) extending radially into the center port and inthis case forming elongated undulations that extend axially through thecenter port. These undulations are present in each concentric circularring-shaped beaded structure such that as one ring-shaped beadedstructure is ablated, the next-presented ring-shaped structure isrevealed, presenting the same geometric pattern, but with an increasedradius.

In FIGS. 1-3 as well as in FIGS. 10-13B, the individual undulations areidentifiable and have a substantially cylindrical shape. However, inpractice, depending upon the scale and layer thickness, such internaltopology can take the form of a dimple pattern 14 as shown in FIGS. 1-3, a corrugation pattern 92 as shown in FIG. 10 , a truncated pyramidalpattern 110 as shown in FIG. 11 , a truncated pyramidal pattern 120 asshown in FIG. 12 , and an irregular pattern 131 as shown in FIGS. 13Aand 13B, all of which may or may not be perceptible to a viewer'sunaided eye. Alternatively, the geometric pattern 14, 92, 110, 120, 131of each ring-shaped concentric beaded structure may take other forms(e.g., an irregular surface) in order to achieve the objective ofincreasing the surface area available for combustion and to ensure thatincreased surface area persists throughout operation of the hybridrocket engine. Any surface features that create the irregular surfaceand increase the surface area of the combustion port (also referred toas the port wall or the flame wall) are considered within the scope ofthe present invention.

In one exemplary embodiment, each fused stacked layer is formed from aseries of fused concentric ring-shaped beads of solidified materialfeaturing a pattern designed to increase surface area of the boundarywall or flame wall, as compared to a smooth construction, and to presentgrooved, protrusion, or contoured patterns. In one embodiment the centerport wall (also referred to as the boundary wall or flame wall) featuresa rifling pattern designed to induce oxidizer vortex flow persistingthroughout the hybrid rocket engine's operation as the fuel is consumed.

In addition to the fused deposition techniques of additive fabrication,as referred to in the cited commonly-owned patent references, there area number of other additive manufacturing methods that can be employed toproduce hybrid rocket fuel grains according to the present invention andusing a formulation of polymer and nanocomposite aluminum additive,without departing from the spirit and scope of the present invention,including: Stereolithography, Selective Laser Sintering, Powder BedPrinting, and Inkjet Head Printing.

For the examples shown in the various Figures described herein, acomposition of the fuel grain is about 95% by mass AcrylonitrileButadiene Styrene (ABS), a thermoplastic possessing combustioncharacteristics desirable for hybrid rocket engine fuel, and about 5%nanocomposite aluminum. Fuel having this structure is available fromseveral sources, as known by those skilled in the art.

With a Young's Modulus of 2.0-2.6 GPa, ABS is 460 times less elasticthan HTPB and 38 times less elastic than paraffin wax, making it anideal material for fabricating a hybrid rocket fuel grain and itscombustion chamber center port.

Ultra-high energetic nano-particle sized aluminum, especially aluminumpowder produced without an aluminum oxide shell and passivated (byencapsulating or ‘capping’ the particle in a polymer shell) for safehandling and use, increases the fuel grain burning rate by as much as50% using only a 5% concentration, compared to a fuel grain fabricatedin ABS with a 25% concentration of standard military grade 44-micronparticle size aluminum.

Referring now to FIGS. 4A-4B, the individual fuel grain sections 10a,10b, 10c, and 10d can be assembled and joined together from multipleseparately fabricated sections to form a complete solid fuel grain 40.In this exemplary embodiment, each solid fuel grain section 10 has aheight, h1, of 23 inches, such that the overall height, h2, of thecomplete solid fuel grain 40 is 92 inches. Furthermore, in thisexemplary embodiment, to ensure proper alignment, the topmost solid fuelgrain 10a has at least one connecting member 100a extending from itslower surface and at least one cavity 102a defined in its lower surfacefor receiving a mating connecting member 104b. Similarly, theintermediate solid fuel grain sections 10b, 10c, each have at least oneconnecting member 100b and 100c, extending from their respective lowersurfaces and one connecting member 104b, 104c, extending from theirrespective upper surfaces, and further each have at least one cavity102b, 102c defined in their respective lower surfaces and at least onecavity 106b, 106c defined in their respective upper surfaces. Finally,the lowermost solid fuel grain section 10d has at least one connectingmember 104b extending from its upper surface and at least one cavity106d defined in its upper surface for receiving a mating connectingmember 100c in the fuel grain section 10c.

Accordingly, when heated above its glass transition temperature butbelow the aluminum powder's ignition temperature, viscous ABS can bespread or sprayed on the upper and lower surfaces to create a strongfusion bond between the sections 10a, 10b, 10c, 10d during assembly. Inthis way, solid fuel grain sections 10a, 10b, 10c, 10d can be readilystacked, aligned, and mated to one another to form the complete solidfuel grain 40.

Referring now to FIG. 5 , after the solid fuel grain sections 10a, 10b,10c, 10d are assembled, the solid fuel grain sections 10a, 10b, 10c, 10dcollectively define a center port 46 through the solid fuel grain 40.The solid fuel grain 40 is preferably wrapped in a film 50 made ofphenol or other suitable thermally resistant material. Placed betweenthe inner wall of a fuel motor case (not shown in FIG. 6 ) and the outersurface of the solid fuel grain, the film 50 acts as an insulation layerto reflect heat and prevent damage to fuel motor cases made from eithermetal or non-metallic materials such as carbon fiber reinforced polymercomposite. Once wrapped in the film 50, the solid fuel grain 40 can beplaced into a motor case of a rocket.

FIG. 6 is a sectional view of an exemplary hybrid rocket engine 70housed within an aeroshell 72 to form a complete hybrid rocket poweredvehicle 70 incorporating the solid fuel grain 40 as described above withrespect to FIGS. 4A, 4B, and 5 . The exemplary hybrid rocket poweredvehicle 70 generally comprises an aeroshell body 72, a nozzle 82 at onedistal end of said aeroshell body 72, and a payload section 74 at anopposite distal end of said aeroshell body 72. Enclosed within the aeroshell body 72 of the hybrid rocket powered vehicle 70 is a hybrid rocketengine including an oxidizer tank 76, a valve 78, a motor case 60, andan oxidizer injector 80 housed typically within a forward cap (notshown) that also houses the ignition system (not shown). The motor case60 houses a pre-combustion chamber (not shown), a post-combustionchamber 64, and the solid fuel grain 40, which as described above iswrapped in insulating film 50.

The solid fuel grain 40 wrapped in a thermal insulating film 50 can be“cartridge loaded” into the motor case 60 of the hybrid rocket engine.Alternatively, the exemplary solid fuel grain 40 wrapped in thermalinsulating film 50 could be wound with a fiber-reinforced polymercomposite to form the motor case without departing from the spirit andscope of the present invention. In another exemplary embodiment, thesolid fuel grain 40 can be inserted into a thermal protection cylinderfabricated from insulating material such as phenolic or cork withoutdeparting from the spirit and scope of the present invention. In yetanother exemplary embodiment, the fuel grain 40 can be formed to embodyeither or both the pre-combustion chamber and the post-combustionchamber 64 without departing from the spirit and scope of the presentinvention.

FIG. 7 is an enlarged sectional view of the motor case 60 of the hybridrocket powered vehicle 70 of FIG. 6 , showing the flame zone within thefuel grain center port 46. As shown, an oxidizer 94 (either a liquid ora gas) is injected into the motor case 60 along a pathway defined by thecenter port 46 of the solid fuel grain 40 and flows within the centerport 46, forming a boundary layer 65 bordered by the center port 46wall. The boundary layer 65 is usually turbulent throughout a largeportion of the length of the center port 46. Within the boundary layer65 is a turbulent diffusion flame zone 66 that extends throughout theentire length of the center port 46 and depending upon thecharacteristics of the solid fuel selected, either causing a phasechange to a gas or entrained liquid droplets of fuel to form.Evaporation from the oxidizer/fuel gas/entrained liquid dropletinterface produces a continuous flow of fuel gas that mixes withoxidizer gas at the flame zone 66 to maintain combustion along theexposed surface area of the center port 46 wall. At steady state, theregression rate of the melt surface and the gas-gas or gas-entrainedliquid droplet interface is the same, and the thickness of the gaseousor entrained liquid layer is constant.

Because the port wall surface pattern 14, 91, 110, 120, 131 exposed tothe flame zone 66 possesses increased surface area compared tocast-molded constructions, the exemplary solid fuel grain 40 causesincreased regression rate and corresponding increased thrust impulsewithout the decreased fuel volumes associated with multi-port designs.Additionally, the undulating wall surface pattern that runs the lengthof the fuel grain port also causes the mixture of fuel gas (or entrainedfuel droplets) and atomized or gaseous oxidizer to continually trip,creating a consistent circular eddy current flow which contributes tomore thorough combustion and a higher Isp.

The continual trip referred to above is a mechanism of motion ofoxidizer and fuel gas through the port. Due to the rough, semi-circularribbed pattern along the port wall (as described elsewhere herein), theoxidizer/fuel gas mixture, as it flows along the boundary with the portwall, will “trip” over the ribs, and create an eddy current. Thistripping mechanism causes the port wall to regress more rapidly,requiring a longer time for the fuel gas mixture to clear the port intothe nozzle; thus contributing to improved combustion and less propellantwaste. This mechanism, together with the much higher surface area thatis created by a ribbed pattern, results in a significantly higher thantypical regression rate as well as higher specific impulse that obtainedwith prior art designs.

Also, unlike the prior art constructions that increase the surface areausing a multi-port architecture (which sacrifices fuel loading), thesolid fuel grain 40 of the present invention allows a smooth burningprocess whereby, as each concentric ring-shaped beaded structure formingeach layer of the fusion stacked layer center port 46 wall is ablated, anew concentric ring-shaped beaded structure, the plurality of whichforms the expanded center port 46 wall, is presented to the flame zone66, as shown in FIGS. 8A-8C, illustrating ablation of the center portwall at three different stages. This burning process continues untileither oxidizer flow is terminated or the solid fuel grain 40 materialis exhausted.

Generally, energetic materials suitable for use in the present inventionare a class of material with high amount of stored chemical energy thatcan be released. Highly energetic materials include ultrafine aluminumpowder, the particle size of which can vary from micron to nanoscale,including particles that are a composite of aluminum and polymer innanoscale. As known by those skilled in the art, generally ananocomposite is a material comprising two or more constituent solids,the size of which measures 100 nanometers (nm) or less. Even though thenano-scale aluminum particle cores are completely encapsulated in apolymer based oligomer coating and thus passivated, there remains thepossibility that this highly energetic pyrophoric material can still bereactive with oxygen or water vapor. As a safety precaution, thenanocomposite aluminum, the ABS thermoplastic, and the compoundedABS-nanocomposite materials are stored in containers designed to storeflammable material, preferably infilled with a non-reactive noble gas atall times prior to their use as feedstock in an additive manufacturingprocess.

In one application, the compounded material is stored within a climatecontrolled environment.

As a further safety measure, after fabrication each fuel grain or fuelgrain section is shrink-wrapped to encase the fuel grain or fuel grainsection in a thin plastic film to prevent atmospheric exposure prior toits use in a hybrid rocket engine. In another embodiment the fuel grainis spray coated with a polymeric material or paint that serves toprevent atmospheric exposure. According to another embodiment the fuelgrain or grain segment is inserted into an air-tight packaging cylinderand a vacuum drawn to remove all air. The packaging cylinder is sealedbefore it is removed from the print bed chamber.

FIG. 9 depicts the coordinate system and orientation of the fuel grainfor use with FIGS. 10-13B

FIG. 10 is a quarter sectional view of the fuel grain section of FIG. 1featuring a concentric ring-shaped corrugation build pattern or fuelgrain 92, a port wall surface pattern 91, and several layers of fusedconcentric beads in cross section 90.

FIG. 11 is a quarter sectional view of the fuel grain section of FIG. 1featuring a concentric ring-shaped truncated pyramidal build pattern orfuel grain 113, a port wall surface pattern 110, and several layers offused concentric beads in cross section 111.

FIG. 12 is a quarter sectional view of the fuel grain section of FIG. 1featuring a concentric ring-shaped rifled truncated pyramidal buildpattern or fuel grain 123, a port wall surface pattern 120 with thebuild and surface patterns staggered layer by layer to form in itsplurality a persistent rifling pattern.

FIG. 13A depicts a top view and FIG. 13B a perspective view showing theport wall surface pattern 131 of the fuel grain section of FIG. 1 .FIGS. 13A and 13B feature a concentric ring-shaped rifled polygonalpattern for fuel grain 132 with each such polygonal build patternstaggered and twisted (i.e., rifled) layer-by-layer to form in itsplurality a persistent rifling pattern.

The embodiments of FIGS. 12 and 13A/13B present a persistent riflingpattern to the oxidizer flowing through the center port 46 to induceaxial flow.

The embodiments of FIGS. 11-13B depict exemplary constructions of ahybrid rocket fuel grain engineered to both increase the amount ofsurface area available for combustion as a means to improve regressionrate, to improve specific impulse, to generate an oxidizer vortex flow,and to reduce fuel waste by inducing oxidizer axial flow within thecenter port 46 (see FIG. 7 ) to allow more time for oxidizer and fuelgases (or oxidizer and entrained liquid droplets) to mix and combustmore thoroughly. Any surface area pattern or topology that furthers oneor more of these objectives, and is sustainable throughout the fuelgrain cross-section (i.e., as one fuel grain layer ablates the next fuelgran layer presents a desirable surface area pattern) is consideredwithin the scope of the present invention.

One of ordinary skill in the art will recognize that additionalembodiments are also possible without departing from the teachings ofthe present invention or the scope of the claims which follow. Thisdetailed description, and particularly the specific details of theexemplary embodiments disclosed herein, is given primarily for clarityof understanding, and no unnecessary limitations are to be understoodtherefrom, for modifications will become obvious to those skilled in theart upon reading this disclosure and may be made without departing fromthe spirit or scope of the claimed invention.

What is claimed is:
 1. A fuel grain for a hybrid rocket, the fuel graincomprising: a plurality of layers of fuel grain material, each layercomprising a plurality of concentric ring-shaped beaded structures ofdifferent radii fused together to form a disc defining a central openingtherein; the plurality of layers stacked and bonded to form acylindrical fuel grain with the central opening of each one of theplurality of layers aligned to form a combustion port extending axiallythrough the fuel grain and bounded by a boundary wall; wherein the fuelgrain material comprises a combustible substance; and an innercircumferential surface of each ring-shaped beaded structure comprisingan irregular surface, such that as a ring-shaped bead forming theboundary wall ablates due to combustion in the combustion port, an innercircumferential wall of an adjacent ring-shaped bead comprising anirregular surface presents to form the boundary wall.
 2. The fuel grainof claim 1 wherein the irregular surface of the boundary wall provides alarger surface area and an increased regression rate of the fuel grainrelative to a fuel grain lacking an irregular surface whilesimultaneously inducing oxidizer/fuel gas flowing through the combustionport to continually trip, thereby creating a consistent eddy current andenabling improved combustion and higher Isp.
 3. The fuel grain of claim1 wherein the plurality of layers of fuel grain material comprises aninner layer forming the boundary wall prior to combustion in thecombustion port, an outer layer forming an outer layer of the fuelgrain, and a plurality of intermediate layers disposed therebetween,wherein progressing from the inner layer to the outer layer anirregularity in the irregular surface is less pronounced.
 4. The fuelgrain of claim 1 wherein the irregular surface comprises projectionsconfigured to form a progressive axial twist through the combustionport, the axial twist for inducing a swirling gaseous flow within thecombustion port.
 5. The fuel grain of claim 4 wherein the progressiveaxial twist comprises a helical grooved rifling pattern of projections.6. The fuel grain of claim 1 wherein the progressive axial twistcomprises a polygonal rifling geometry.
 7. The fuel grain of claim 1produced by an additive manufacturing process.
 8. The fuel grain ofclaim 1 wherein the grain material comprises Acrylonitrile ButadieneStyrene (ABS) thermoplastic and a plurality of micron scale or nanoscaleelemental aluminum particles or a plurality of nanoscale elementalaluminum core particles capped with an oligomer polymer.
 9. The fuelgrain of claim 1 wherein the grain material comprises AcrylonitrileButadiene Styrene (ABS) thermoplastic by mass ranging from 80% to 95%and aluminum powder by mass correspondingly ranging from 20% to 5%, theparticle size of which can vary from 15 nanometers to 44 microns. 10.The fuel grain of claim 1 further comprising a thermally insulatingmaterial encasing the fuel grain.
 11. The fuel grain of claim 1 theirregular surface comprising one or more of a plurality of ribs, aplurality of undulations, a plurality of protrusions and recesses, aplurality of depressions.
 12. The fuel grain of claim 1 the irregularsurface comprising one or more of a corrugation pattern, a truncatedpyramidal pattern, a rifled truncated pyramidal pattern, or a rifledpolygonal pattern.
 13. The fuel grain of claim 1 wherein a shape of thecombustion port comprises a circular shape, an oval shape, an ellipticalshape, a polygonal shape, a quatrefoil shape, a star shape, or anirregular shape.
 14. The fuel grain of claim 1 wherein the fuel graindefines an outer diameter of 19.0 inches and the combustion port has aninitial diameter of 4 inches prior to consumption of fuel grain materialduring combustion.
 15. A plurality of fuel grain segments each accordingto claim 1 further comprising ABS material between a surface of a firstfuel grain segment and an abutting surface of a second fuel grainsegment thereby creating a fusion bond between the first and second fuelgrain segments.
 16. The fuel grain of claim 1 wherein a composition ofthe fuel grain material of each one of the concentric ring-shaped beadedstructures is substantially uniform.
 17. The fuel grain of claim 1wherein the combustible substance comprises a formulation ofthermoplastic and passivated nanoscale metallic material.
 18. The fuelgrain of claim 1 wherein the irregular surface of each ring-shapedbeaded structure comprises a sustaining internal topological pattern aseach ring-shaped beaded structure ablates and another ring-shaped beadedstructure is revealed due to combustion and ablation to the combustionport wall.
 19. The fuel grain of claim 1, wherein the irregular surfaceof each ring-shaped beaded structures forms a sustaining rifling patternor geometry for both increasing the surface area of the combustion portwall and for generating a vortex flow of oxidizer and fuel gas flowingthrough the combustion port.
 20. The fuel grain of claim 1 wherein amaterial of each ring-shaped beaded structure comprises a solidifiedmaterial, further comprising a polymer or a solidified polymer-metalblend formulation suitable for combusting in a hybrid rocket engine. 21.The fuel grain of claim 1 wherein a material of each ring-shaped beadedstructure comprises a blend of Acrylonitrile Butadiene Styrene (ABS) andaluminum powder.
 22. The fuel grain of claim 1 wherein the combustionport defines a polygonal shape in a cross section with an orientation ofeach layer adjusted to create a progressive helical twist axiallythrough the combustion port, forming a rifling pattern to induce aswirling oxidizer/fuel gaseous flow within the center combustion port.23. A fuel grain for a hybrid rocket, the fuel grain comprising: a firstfuel grain section comprising: a first plurality of concentricring-shaped beads of different radii fused together to form a firstdisc, the first disc defining a first combustion port; an innercircumferential surface of each of the first plurality of circularring-shaped beads comprising an irregular surface, such that as aring-shaped bead forming a first combustion port wall ablates due tocombustion in a combustion port, an inner circumferential surface of anadjacent ring-shaped bead comprising an irregular surface presents toform the first combustion port wall; a material of the first fuel grainsection comprising a combustible substance; a second fuel grain sectioncomprising: a second plurality of concentric ring-shaped beads ofdifferent radii fused together to form a second disc, the second discdefining a second combustion port; an inner circumferential surface ofeach of the second plurality of circular ring-shaped beads comprising anirregular surface, such that as a ring-shaped bead forming a secondcombustion port ablates due to combustion in a combustion port, an innercircumferential wall of an adjacent ring-shaped bead comprising anirregular surface presents to form the second combustion port wall; amaterial of the second fuel grain section comprising a combustiblesubstance; and the first and second fuel grain sections bonded togetherand the first and second combustion ports aligned to form the fuelgrain.
 24. The fuel grain of claim 23 further comprising a firstconnecting member in a lower surface of the first fuel grain section formating with a second connecting member in an upper surface of the secondfuel grain section.
 25. A hybrid rocket engine comprising: a fuel grainfurther comprising: a plurality of layers of fuel grain material eachlayer comprising a plurality of concentric ring-shaped beaded structuresof different radii fused together to form a disc, the disc defining acentral opening; the plurality of layers stacked and bonded to form acylindrical fuel grain such that the central opening of each one of theplurality of layers is aligned to form a combustion port extendingaxially through the fuel grain and bounded by a boundary wall; whereinthe fuel grain material includes at least one combustible substance; aninner circumferential surface of each ring-shaped beaded structurecomprising an irregular surface, such that as a ring-shaped bead formingthe boundary wall ablates due to combustion in the combustion port, aninner circumferential wall of an adjacent ring-shaped bead comprising anirregular surface presents to form the boundary wall; an oxidizersource, the oxidizer for flowing through the combustion port duringengine operation; a valve for controlling flow of oxidizer through thecombustion port; a nozzle in fluid communication with the combustionport; and a shell for housing the fuel grain, the oxidizer source, andthe valve, the nozzle extending from the shell.
 26. The hybrid rocketengine of claim 25 further comprising an insulating film surrounding thefuel grain.
 27. A fuel grain for a hybrid rocket, the fuel graincomprising: multiple beads of fuel grain material, in which the beadsare fused together to form a generally cylindrical fuel grain defining acombustion port extending axially through the generally cylindrical fuelgrain, in which the combustion port is bounded by a boundary wall, inwhich the fuel grain material comprises a combustible substance, and inwhich the boundary wall is formed of a subset of the beads ofcombustible fuel grain material, and in which the boundary wall isconfigured to induce an eddy current in a fluid flowing through thecombustion port.
 28. The fuel grain of claim 27, in which the fuel grainis configured such that when the boundary wall ablates due to combustionin the combustion port, a new surface of fuel grain material is exposedto the combustion port.
 29. The fuel grain of claim 27, in which theboundary wall presents a larger surface area to the combustion portrelative to a boundary wall that is not formed of beads of fuel grainmaterial.
 30. The fuel grain of claim 27, in which the boundary wall istextured with one or more of ribs, dimples, undulations, protrusions, ordepressions.
 31. The fuel grain of claim 27, in which when viewed alonga longitudinal axis of the combustion port, the boundary wall definesalternating protrusions and depressions.
 32. The fuel grain of claim 27,in which the fuel grain is fabricated in an additive manufacturingprocess.
 33. The fuel grain of claim 27, in which the fuel grainmaterial comprises an Acrylonitrile Butadiene Styrene (ABS)thermoplastic.
 34. The fuel grain of claim 27, in which the fuel grainmaterial comprises a mixture of a hybrid rocket fuel material and ananoscale metallic material.
 35. The fuel grain of claim 34, in whichthe fuel grain material comprises between 80% and 95% by mass of thehybrid rocket fuel material and between 5% and 20% by mass of thenanoscale metallic material.
 36. The fuel grain of claim 27, comprisinga thermally insulating material encasing the fuel grain.
 37. A fuelgrain assembly comprising: multiple of the fuel grains of claim 28, inwhich an end of fuel grain is bonded to an end of an adjacent fuel grainto form an elongated, generally cylindrical fuel grain assembly, and inwhich the combustion ports of the multiple fuel grains are aligned todefine an elongated combustion port of the fuel grain assembly.
 38. Ahybrid rocket engine comprising: the fuel grain of claim 28; an oxidizersource configured to provide a flow of an oxidizer through thecombustion port during operation of the hybrid rocket engine; a valveconfigured to control the flow of the oxidizer through the combustionport; a nozzle in fluid communication with the combustion port; and acasing, in which the fuel grain, the oxidizer source, and the valve arehoused within the casing, and in which the nozzle extends beyond an endof the casing.
 39. A fuel grain for a hybrid rocket, the fuel graincomprising: a generally cylindrical body formed of a fuel grain materialcomprising a combustible substance, in which a combustion port extendsaxially through the body, in which the combustion port is bounded by aboundary wall, in which the boundary wall is formed of the fuel grainmaterial, and in which the boundary wall is configured to induce an eddycurrent in the combustion port in a fluid flowing through the combustionport, and in which the fuel grain is configured such that when theboundary wall ablates due to combustion in the combustion port, a newsurface of fuel grain material is exposed to the combustion port. 40.The fuel grain of claim 39, in which the body comprises beads of thefuel grain material, in which the beads are fused together to form thebody.
 41. The fuel grain of claim 39, in which the boundary wall istextured with one or more of ribs, dimples, undulations, protrusions, ordepressions.
 42. The fuel grain of claim 41, in which the boundary wallpresents a larger surface area to the combustion port relative to aboundary wall that is not textured.
 43. The fuel grain of claim 39, inwhich when viewed along a longitudinal axis of the combustion port, theboundary wall defines alternating protrusions and depressions.
 44. Thefuel grain of claim 39, in which the fuel grain is fabricated in anadditive manufacturing process.
 45. The fuel grain of claim 39, in whichthe fuel grain material comprises an Acrylonitrile Butadiene Styrene(ABS) thermoplastic.
 46. The fuel grain of claim 39, in which the fuelgrain material comprises a mixture of a hybrid rocket fuel material anda nanoscale metallic material.
 47. The fuel grain of claim 46, in whichthe fuel grain material comprises between 80% and 95% by mass of thehybrid rocket fuel material and between 5% and 20% by mass of thenanoscale metallic material.
 48. The fuel grain of claim 39, comprisinga thermally insulating material encasing the fuel grain.
 49. A fuelgrain assembly comprising: multiple of the fuel grains of claim 39, inwhich an end of fuel grain is bonded to an end of an adjacent fuel grainto form an elongated, generally cylindrical fuel grain assembly, and inwhich the combustion ports of the multiple fuel grains are aligned todefine an elongated combustion port of the fuel grain assembly.
 50. Ahybrid rocket engine comprising: the fuel grain of claim 39; an oxidizersource configured to provide a flow of an oxidizer through thecombustion port during operation of the hybrid rocket engine; a valveconfigured to control the flow of the oxidizer through the combustionport; a nozzle in fluid communication with the combustion port; and acasing, in which the fuel grain, the oxidizer source, and the valve arehoused within the casing, and in which the nozzle extends beyond an endof the casing.